1. Field of the Invention
The present invention relates generally to gas turbine engines, and more specifically to turbine blade cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
One method of improving the efficiency of a gas turbine engine is to increase the temperature of the hot gas stream that passes through the turbine. In order to allow for a higher gas temperature in the turbine, one way designers meet this challenge is to provide more effective blade cooling in order that the blade materials can withstand the higher temperature.
Turbine blades are therefore cooled by passing a cooling fluid such as compressed air through serpentine passageways in the blade. Cooling air is also discharge into the gas stream through cooling holes strategically placed to provide an air cushion on the hottest surfaces of the blade. Examples of cooling methods for turbine blades include convection cooling and impingement cooling in which the cooling fluid passes through the inside of the turbine blade, and film cooling in which the cooling fluid is ejected to the outside surface of the turbine blade to form a film of cooling fluid.
Squealer tips have been used on the tips of turbine blades to provide a seal between the rotating turbine blade and the stationary blade outer air seal (BOAS). Increased engine efficiency is obtained when the gap between the tip and the turbine shroud is minimized. The tip clearance is limited by the differential thermal expansion and contraction between the blade and the turbine shroud. If rubbing occurs, the effects will be minimal because of the low surface area exposed to the rubbing due to the squealer tips. Leakage of the hot gas flow through the gap formed between the blade tip and the turbine shroud decreases the efficiency of the engine, and also allows for the blade tip and blade outer surface to be exposed to the hot gas flow that can damage the blade and tip. The squealer tip is typically of small thickness and particularly susceptible to high temperature oxidation and other damage due to over-heating. The blade tip is particularly difficult to cool since it is located directly adjacent to the turbine shroud, and the hot combustion gas flow passes through the tip gap.
High temperature turbine blade tip section heat load is a function of the blade tip leakage flow. A high leakage flow will induce a high heat load onto the blade tip section. Therefore, the blade tip section sealing and cooling must be addressed as a single problem. Traditionally, a typical turbine blade tip includes a squealer tip rail which extends around the perimeter of the airfoil flush with the airfoil wall such that an inner squealer pocket is formed. The main purpose of incorporating a squealer tip in a blade design is to reduce the blade tip leakage and also to provide for rubbing capability for the blade. FIG. 1 shows a squealer tip cooling arrangement. Film cooling holes are formed along the airfoil pressure side tip section, and extend from a leading edge to a trailing edge in order to provide for edge cooling of the blade pressure side squealer tip. In addition, convective cooling holes are also formed along the tip rail at an inner portion of the squealer pocket in order to provide additional cooling for the squealer tip rail. Secondary hot gas flow migration around the blade tip section is also shown in FIG. 1. the blade includes a pressure side 110, a squealer tip 134 forming a squealer pocket 124, cooling holes along the pressure side airfoil surface, and cooling holes 122 adjacent to the sides of the squealer tip 134.
U.S. Pat. No. 6,994,514 B2 issued to Soechting et al on Feb. 7, 2006 shown in FIG. 2 shows a TURBINE BLADE AND GAS TURBINE with a cooling concept for the blade suction side tip rail in which the blade includes a pressure side 236, a suction side 235, a squealer tip 237, film cooling holes 238 near to suction side, and film cooling holes 239 near the pressure side for the blade. The suction side blade tip rail 237 is subject to heating due to the hot gas flow over the blade tip from three exposed sides, cooling of the suction side squealer tip rail 237 by means of discharge row of film cooling holes 239 along the blade pressure side peripheral and at the bottom of the squealer floor becomes insufficient. This is primary due to the combination of tip rail geometry and the interaction of hot gas secondary flow mixing, whereby the effectiveness induced by the pressure side film cooling and tip section convective cooling holes is very limited.
FIG. 3 is from the U.S. Pat. No. 6,527,514 B2 issued to Roeloffs on Mar. 4, 2003 entitled TURBINE BLADE WITH RUB TOLERANT COOLING CONSTRUCTION and shows a turbine blade with a pressure side 302, a suction side 303, a tip cap 304 having an inner surface 314, a blade hollow space 305, a pressure side tip crown 307, a suction side tip crown 308, a pressure side cooling passage 325 opening onto film cooling holes 310 on the pressure side surface of the blade, and a cooling passage 315 extending in a first portion 317 from the hollow space 305 through the tip cap 304 to an exit hole opening into a cavity 316 and then through an exit hole 311 opening onto the suction side tip squealer 308. A tip pocket 309 is formed between the two squealer tips 307 and 308.
FIG. 4 is from the U.S. Pat. No. 6,602,052 B2 issued to Liang on Aug. 5, 2003 and shows an AIRFOIL TIP SQUEALER COOLING CONSTRUCTION in which a turbine blade with a pressure side 402 and suction side 403, a blade tip cap 404 with a pressure side squealer tip 407 and a suction side squealer tip 408, a tip pocket 409 formed between the two squealer tips 407 and 408, film cooling holes 418 opening onto the pressure side airfoil surface, and film cooling holes 414 adjacent to the suction side squealer tip 408.
FIG. 5 is form the U.S. Pat. No. 6,059,530 issued to Lee on May 9, 2000 entitled TWIN RIB TURBINE BLADE and shows a turbine blade with a pressure side 528 and a suction side 530, a first squealer tip 550 and a second squealer tip 552, a tip channel 554 formed between the two tips 550 and 552, an internal flow channel or chamber 540, and two film cooling holes 562 to supply cooling air to the pressure side of the first tip 550. Cooling air is also discharged into the tip channel 554 for mixing with the combustion gases to further decrease the temperature of the gases for cooling both tip ribs and their inboard sides.
FIG. 6 shows the U.S. Pat. No. 6,991,430 B2 issued to Stec et al on Jan. 31, 2006 entitled TURBINE BLADE WITH RECESSED SQUEALER TIP AND SHELF with a turbine blade having a pressure side 624, a suction side 626, a continuous tip squealer wall 662 extending around the tip of the blade and forming a tip cavity 640, and a recessed tip wall portion 645 recessed inboard from the pressure side of the airfoil wall forming a tip shelf 647 there between. A plurality of film cooling shelf holes 652 in the tip cap 622 supply cooling air to the recessed tip wall 645, and a plurality of film cooling holes 646 supply cooling air to the tip cavity 640.
It is an object of the present invention to provide for turbine blade of a gas turbine engine with improved tip cooling.
It is another object of the present invention to provide for a turbine blade tip with improved sealing between the tip and the turbine shroud.